Fluid propulsion devices achieve thrust by imparting momentum to a fluid called the propellant. An air-breathing engine, as the name implies, uses the atmosphere for most of its propellant. The gas turbine produces high-temperature gas which may be used either to generate power for a propeller, fan, generator or other mechanical apparatus or to develop thrust directly by expansion and acceleration of the hot gas in a nozzle. In any case, an air breathing engine continuously draws air from the atmosphere, compresses it, adds energy in the form of heat, and then expands it in order to convert the added energy to shaft work or jet kinetic energy. Thus, in addition to acting as propellant, the air acts as the working fluid in a thermodynamic process in which a fraction of the energy is made available for propulsive purposes or work.
Typically turbofan engines include at least two air streams. All air utilized by the engine initially passes through a fan, and then it is split into the two air streams. The inner air stream is referred to as core air and passes into the compressor portion of the engine, where it is compressed. This air then is fed to the combustor portion of the engine where it is mixed with fuel and the fuel is combusted. The combustion gases are then expanded through the turbine portion of the engine, which extracts energy from the hot combustion gases, the extracted energy being used to run the compressor, the fan and other accessory systems. The remaining hot gases then flow into the exhaust portion of the engine, which may be used to produce thrust for forward motion to the aircraft.
In the turbofan engine shown in FIG. 1, the flow of the air is generally axial. The engine direction along the axis is generally defined using the terms “upstream” and “downstream” generally which refer to a position in a jet engine in relation to the ambient air inlet and the engine exhaust at the back of the engine. For example, the inlet fan is upstream of the combustion chamber. Likewise, the terms “fore” and “aft” generally refer to a position in relation to the ambient air inlet and the engine exhaust nozzle. Additionally, outward/outboard and inward/inboard refer to the radial direction. For example, the bypass duct is outboard the core duct. The ducts are generally circular and co-axial with each other.
As ambient inlet airflow 12 enters inlet fan duct 14 of turbofan engine 10, the air passes through the guide vanes 15, by fan spinner 16, and through fan rotor (fan blade) 42. The airflow 12 is split into primary (core) flow stream 28 and bypass flow stream 30 by upstream splitter 24 and downstream splitter 25.
In FIG. 2, the bypass flow stream 30 along with the core/primary flow stream 28 is shown, the bypass stream 30 being outboard of the core stream 28. The inward portion of the bypass steam 30 and the outward portion of the core stream 28 are partially defined by the splitters upstream of the compressor 26. The fan 42 has a plurality of fan blades.
As shown in FIG. 1 the fan blade 42 is rotating about the engine axis into the page, therefor the low pressure side of the blade 42 is shown, the high pressure side being on the opposite side. The primary flow (core) stream 28 flows through compressor 26 that compresses the air to a higher pressure. The core flow stream 28 is then mixed with fuel in combustion chamber 35 and the mixture is ignited and burned. The resultant combustion products flow through turbines 38 that extract energy from the combustion gases to turn fan rotor 42, compressor 26 and any shaft work by way of turbine shaft 40. The gases, passing through the exhaust cone, expand through an exhaust nozzle 43 to produce thrust. Primary flow stream 28 leaves the engine at a higher velocity than when it entered. Bypass flow stream 30 flows through fan rotor 42, flows by bypass duct outer wall 27 (an annular duct concentric with the core engine), flows through fan discharge outlet, and is expanded through an exhaust nozzle to produce additional thrust. Turbofan engine 10 has a generally longitudinally extending centerline represented by engine axis 40.
Current compressor design relies on conventional airfoils, borrowed from aircraft wing theory, disposed in an annular duct. The design of both low- and high-subsonic airfoils in modern, axial-flow compressors has remained essentially the same since the mid-1980's when Hobbs and Weingold published their work on what is now known as the controlled-diffusion airfoil. Improvements in the performance of low- and high-subsonic axial-flow compressors having controlled-diffusion airfoils have been realized by optimizing the solidity, aspect-ratio, and three dimensional stacking of these airfoil sections leading to loss-reduction and increasing the operable incidence range. Further performance improvements have been realized by minimizing the blade-tip and stator shroud clearances, leakage paths, and bleed flows that can disrupt the flow in the blade rows. Finally, improvements to simulation tools have allowed designers to more accurately set the stage-matching of blade rows.
As the above approaches to improve airfoil performance have matured, it has become increasingly difficult to attain further improvements for modern, axial-flow compressors. State-of-the-art compressor technology has continued to show an asymptotic trend in performance.
Despite the design limitations of conventional compressors, the demand to increase compressor performance, and thereby reduce engine fuel consumption, remains high. There remains a need to minimize aerodynamic losses and increase the incidence-range of axial-flow compressor blading to reduce specific fuel consumption and improve the operating range of the attendant blading.
The purpose of the boundary layer control, also known as BLC, is to affect the flow by influencing the structure of the boundary layer, in order to increase the efficiency, the loading and the stage pressure ratio of turbo engines, and of design performance of isolated airfoils and bodies. The main advantage of boundary layer control discussed herein is to prevent or delay boundary layer separation and thereby increase the allowable blade or airfoil loading and range of angles of attack.
The application of flow control to axial-compressor blade/vane row design allows an increase in blade/vane loading levels and a broadened operating range. In particular, the increase in loading of blade/vane rows results in fewer blades/vanes and/or stages used for the same magnitude of flow turning and compressor pressure ratio respectively. The reduction in parts count yields a reduction in wetted area, reduced maintenance and reduced total weight of the compressor.
Fluid dynamics have been considered for a substantial period of time. One of the arts in which substantial and powerful thought has been devoted to is that of compressors and turbomachines. One of the most important areas driving such research is aeronautics and astronautics for both the commercial interests of high speed transportation and military interests for defense and the exploration of space. Some important issues with respect to the advance of compressors and turbomachines is the attainable pressure ratio and the efficiency of the machines.
Reissue U.S. Pat. No. 23,108 to E. A. Stalker discloses the provision of slots located well rearward on the blade to increase the effectiveness of the blade. This is taught in order to control the boundary layer on the blades of blowers and compressors to better enable the machine to run at lower than optimal speeds.
J. R. Irwin, U.S. Pat. No. 2,720,356 imposes continuous boundary layer control for compressors by moving the boundary layer through porous surfaces. The teaching recommends to then reintroduce the viscous interactive flow to the main flow of the compressor at a later stage.
U.S. Pat. No. 2,749,025 to Stalker focuses primarily on providing blades of later stages in a compressor with progressively larger radii rounded leading edges. This reduces losses associated with the flow angle into these blades which would normally be experienced at below optimum speeds. The substantially semi-circular nose cross-section is professed to be able to smooth the flow and avoid burbling when the approach vectors are far from optimum. A further step to assist the machine in these conditions is to remove the boundary layer in this area.
U.S. Pat. No. 3,694,102 to Conrad teaches use of suction slots in stator blades to prevent separation of the boundary layer in supersonic blading. Conrad, however, fails to recognize the benefit of removing the boundary layer permanently from a compressor. This is evidenced by equating bleeding of the boundary layer to atmosphere to reintroducing the boundary layer into the compressor at another stage.
U.S. Pat. No. 3,993,414 to Meauze discloses an axial supersonic compressor comprising a casing and a hub rotating in the casing and carrying blades. On each of the suction surfaces of the blades is formed a zone in which the curvative changes and which corresponds to a supersonic shock wave. A channel formed in each blade and opening in the zone is connected to a boundary layer aspiration means.
U.S. Pat. No. 3,897,168 to Amos and U.S. Pat. No. 4,595,339 to Naudet both disclose the recapture of energy from a withdrawn boundary layer to avoid losses.
U.S. Pat. No. 3,385,509 to Gamier discloses an engine with counter-rotating compressor blades and counter-rotating turbine blades. Nozzle flow area of the turbines is adjusted to control the boundary layer by either moving the stators or by blowing through slots in their surfaces. Gamier is silent however on removing the boundary layer from the flow permanently.